A manned Mars program made up of six missions is evaluated to determine the impact of using lunar liquid oxygen (LOX) as a propellant. Two departure and return nodes, low Earth orbit and low lunar orbit, are considered, as well as two return vehicle configurations, a full 70,000-kg vehicle and a 6800-kg capsule. The cost of lunar LOX delivered to orbit is expressed as a ratio of Earth launch cost

Tucker, William B.

Young, Roy M., Jr.

English

NASA Technical Reports Server

453
IMPACT OF LUNAR OXYGEN PRODUCTION
ON DIRECT MANNED MARS MISSIONS
Roy M. Young Jl'. 1 and William B. Tucker 2 _ 9 3 -
SRS Technologies
990 Explorer Bk,d.
Huntsville AZ 35806
LU
A manned Mars program made up of six mi._'o_zs is elvMuated to determine the irnlx_'t of using lum, r
liquid oxygen (LOX) as a propellant. _ de[torture aml return nodes, lou, Earth orbit and Iou, lunar
orbit, are considered, as u._ll as tu_) return t_ehicle configurations, a full 70,(X)O-kg tehicle and a (_800-
kg capsule. The cost of lunar LOX deli_c, red to orbit is e._qTmessed as a ratio of Fwlrth launch cost.
LIST OF ACRONYMS
TLI Trans Lunar Injection
LOI Lunar orbit insertion
TMI Trans-Mars injection
MOI Mars orbit insertion
TEl Trans-Earth injection
EOI Earth orbit insertion
EEC Earth departure and Earth return capsule vehicle
EEF Earth departure and Earth return full vehicle
LEC Lunar departure and Earth return capsule vehicle
LEF Lunar departure and Earth return full vehicle
LLC Lunar departure and lunar return capsule vehicle
LLF Lunar departure and lunar return full vehicle
ECCV Earth crew capture vehicle
MTV Mars transfer vehicle
ETV Earth transfer vehicle
LEO Low Earth orbit
LOX Liquid oxygen
LH 2 Liquid hydrogen
LLO Low lunar orbit
INTRODUCTION
As part of the SRS Manned Mars Mission and Program Analysis
Study conducted for Marshall Space Flight Center, a series of
manned Mars programs was investigated, each culminating with
a manned base on Mars. Initial guidelines and a.csumptions for the
study specify the 1999-2035 timeframe and that each program
would includc an exploratory, outpost, and base pha.se. The first
three manned missions of each pr(moram were considered site
survey missions. The outpost phase was reached wqaen a portion
of the crew had the option of remaining at Mars between
mi._sions. The base phase was accomplished when there was a
permanently manned facility on the surface of Mars.
t Now at Mail Stop EO24, Marshall Space Flight Center, Huntsville
AL 35812
2Member of The Technical .Staff, Nichols Research Corp., 4040 S. Memorial
Parkway, Huntsville AL 35806
KS Fig. 1 shows, each of the five example programs emphasized
different objectives. The Early (or Reference) Program takes
ath_antage of the cyclic nature of Earth-Mars oppositions (see
Fig. 2) to accomplish the shorter exploratory missions near the
oppositions when Mars is at perihelion, and represents a fairly
vigorous program to reach milestones as ._)n as ix)ssible. The
Phobos/Deimos Program has initial landings on the surface of the
martian meatus and later excursions to the surface of Mars. The
Split/Sprint Program uses .split missions with cargo vehicles _nt
on long, low-energy trajectories, and the lighter manned mtxfules
sent on shorter, high-energy trajectories. The Mid-Range Program
delays the start of a manned Mars program while additional
emphasis is placed on low Earth orbit (LEO) activities. The Later
Program delays the missions for one cycle of oppositions,
approximately 15years, and includes a lunar emphasis program
for the years 1999-2018. This pal-mr investigates the lXJ_sible
benefits of a lunar liquid oxygen (IOX) propellant plant on the
Later Manned Mars Program.
Studies on the impact of lunar LOX production on Mars
migsions have been conducted by several authors (Keaton, 1985;
Babb and Stump, 1986; Ctrrdell and Wagner, 1987). These
studies have all been based on conjunction class "minimum
energy" missions. Conjunction missions use low-energy near-
Hohmann transfer trajectories, which cause flight times to Mars
to be on the order of eight months each way. Waiting at Ma_
for a near-Hohmarm return opportunity reqmres stay times at Mars
of 300 to 550 days. Thus the total mission time for a coniunction
class mission is 2.5 to 3 years. Conjunction class missions provide
the most efficient use of propellants when translx_rting large
amounts of material to and from Mars..'_.nsitivity studies (Dat_,
1986) have shown that the conjunction cla.ss mission can deliver
four times the payload on a round-trip mission or six times the
one-way payload from Earth orbit to Mars orbit than the
opposition class mission (with the .same initial mass in LEO and
propulsion technology). The dimth-antage of the conjunction cla.ss
mission lies in the long total trip time that can Ix" debilitating
on the crew and increase propellant boil-off losses. For this reason
the opposition class mission is important in planning Mars
programs. The oplx_sition clas,s trajectot 7 is characterized by 30-
90-day stay times at Mars and total trip times from 450 to
730 da)._. This paper considers opposition class missinns for the
manned exploratory and early outpost phases of the program and
https://ntrs.nasa.gov/search.jsp?R=19930004796 2020-03-17T10:16:08+00:00Z
454 2nd Conference on Lunar Bases and Space Activities
Early
Mid-Range
Later
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Fig. 1. Example of the manned Mars mission programs.
2O22
2023 200_2005
'o"q/L Y<"
III --"-i _ :'_".'._- _ V- |111 M,_ P_
201_ 2018
Fig. 2. Earth-Mars oppositions.
conjunction class missions for the later outpost and base phases
of the program. Table 1 details the missions required to accom-
plish the program with the payloads carried on each mission.
COMPARATIVE SCENARIOS
The Lax propellant required for the various missions in the
Later Program was assumed to be produced on either the Earth
or the Moon. In either case a 500-km circular space station orbit
is used as a staging area for assembly of the Mars transfer vehicle
(MTV). If the Lax is produced on the Earth the trans-Mars
injection (TMI) burn is initiated from the LEO staging orbit. If
the Lax is produced at a lunar propellant plant the MTV is loaded
at the staging area with all the liquid hydrogen (LH2) required
for the mission, including the lunar operations, and with enough
Lax to initiate the translunar injection (TLI) and lunar orbit in-
sertion (LOI) maneuvers. The AVs for these burns are 3.155 kin/
sac and 0.975 kin/sac respectively (Babb and Stump, 1986). The
resulting lunar orbit is a 500-km circular orbit. This study assumes
that the lunar Lax is available in the same orbit. The MTV is then
TABLE 1. Mission characteristics for the "Later" Program.
Mission (_ition Trip Tune Stay Time
Number Mission Class Year in Days in Days Payloads
l Manned opposition 2018 450 60 130,000 kg
2 Manned opposition 2020 550 60 130,000 kg
3 Manned opposition 2022 730 90 130,000 kg
4 Manned opposition 2025 730 90 178,000 kg
5a Manned opposition 2027 730 90 130,000 kg
5b Cargo contraction 2026 220 NA 100,000 kg
6 Manned conjunction 2029 1000 500 230,000 kg
Young and Tucker: Lunarpropellantproduction 455
tanked up with the lunar LOX and the TMI operation is initiated.
The optimum time for lunar departure is at a full Moon when
the velocity vectors for the Earth and Moon are in the same di-
rection. All departures from and returns to the Moon were there-
fore constrained to occur at the time of the nearest full Moon.
The MTV used in this analysis is a three-stage system with each
stage's propulsion system having an lap of 482 sec. The propellant
used is a mixture of LOX and LH 2 with a mixture ratio of 6:1.
The first stage performs the TMI burn and is then expended. The
second stage performs a 25 m/see midcourse correction, the Mars
orbital insertion (MOI), and trans-Earth injection (TEI) maneu-
vers. The MOI maneuver is accomplished with an aerobrake pro-
vided the C3 at Mars is less then 50km/sec, otherwise a
propulsive bum is required to reduce the C3 to 50 km/sec. A
24-hr 500-km periapsis highly elliptical parking orbit around Mars
was assumed to accommodate Mars landing operations and
minimize departure AV requirements. After the appropriate stay
time at Mars, the second stage is ignited to perform the TEl.
Following this bum the second stage is expended and the Earth
transfer vehicle (ETV) utilizes the third stage for a midcourse
burn of 25 m/see.
At this point in the analysis two return vehicle configurations
are investigated. A full module return brings back the entire ETV
with a mass of 70,000 kg. A capsule return expends the ETV and
crew return is completed in a small Apollo-type Earth crew
capture vehicle (ECCV) with a mass of only 6800 kg. Two return
nodes are also investigated in the study: return to a 5OO-km
circular low lunar orbit (LLO) or return to a 24-hr 5OO-km
perigee highly elliptical Earth orbit. If the ETV or ECCV is
returned to Earth orbit, an aerobrake is assumed for the Earth
orbit insertion (EOl) maneuver if the C3 at Earth arrival is less
then 25 km/sec. If the C3 is greater than 25 km/sec the EOI
maneuver is performed propulsively by the third stage. In both
the full return and capsule return to the lunar orbit, the LOI is
performed by the third stage. &s was the case with lunar
departure, return to the Moon is accomplished at a full Moon to
minimize the AV requirements. Figure 3 depicts the various
options available for departure and return nodes.
Near Earth Departure Phase
Earth.Mars
Hellocentrl
Phase
Near Mars Operations Phase Mars-Earth
Hellocent rio
Phase
Mars-Earth
Heliocentric
Phase
Near Earth Return Phase
EIrth Ca ureOp.rN_
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Eirth Atre, os_m B
Lunlr R(ml¢ovl Ol_mllon
letup_ t_
; .:;
'_..dszc=_ " :
1 t_ow_._l o o
Fig. 3. Options available for departure, arrival, and return nodes.
456 2nd Conference on Lunar Bases and Space Activities
RESULTS AND CONCLUSIONS
The total trip time, stay time at Mars, and payload masses
required tor each mission are inputs to the SRS trajectory analysis
code. Using Lambert's theorem to calculate AV requirements, the
p_l determines the trajectory, fuel requirements, and vehicle
size to minimize the mass in LEO or LLO, depending on departure
point. If a lunar departure is planned, the fuel requirements for
the TLI and LOI are also calculated and charged to the mass
required at the LEO staging orbit. The results of these calculations
are presented in Fig. 4. Note that the mass of the payload is not
included in the total mass since it is constant for each mission.
As can be seen from Fig. 4 there is a wide variation in the total
mass required to accomplish each mission in the later Program.
This is due to the different payloads carried (see Table 1 ) and
the oppositions variations in Earth/Mars (see Fig. 2). As would be
expected, the capsule returns are the best options due to the
much smaller mass of the ECCV. In addition, the initial mass
required in LEO for lunar departure, including the extra
propellants for TLI and LOI, are generally less than the mass in
LEO required for direct missions from the Earth to Mars. The
lunar departure-Earth retm'n options provide the best lx'rfor-
mance since the AV requirements for TMI are smaller when
departing from lunar orbit and the AV requirements at Earth
return are smaller due to the higher orbital velocity in LEO. The
lunar departure-lunar return options are generally the worst, due
Total
Mass
Payload
(rot)
5000
4OOO
3000
2OO0
1000
0
2018
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_mt_m_J_l _ Lunar LOX F
EEF LEF LLF EEC LEC LLC
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2020 _ I I
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BI LEoLOXI:
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Mass 2222222_///'__E___ _/'///'_
(mr) 20oo _
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o
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Lunar LOX
LEF LLF EEC LEC LLC
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Total
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Payload
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Total
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2000
1000
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o
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EEF - Earth - Earth - Full Vehicle
LEF - Lunar- Earth - Full Vehicle
LLF - Lunar - Lunar - Full Vehicle
EEC - Earth - Earth - Capsule
LEC - Lunar - Earth - Capsule
LLC - Lunar - Lunar - Capsule
Fig. 4. Mass requirements for missions in the "later" Program ( ! mt = 10OO kg).
Young and Tucker: Lunarpropellant production 457
primarily to the large AV maneuver required to place the vehicle
in LLO upon return. Figure 5 shows AV requirements for the
different missions and options.
Determining the cost savings incurred by using lunar ILOX is
a difficult task. The total cost of LOX in LEO is dominated by
the Earth surface to LEO transportation cost, with mantffacturing
cost being negligible. The transportation cost t)f lunar LOX should
be less, based on the smaller gravitational field; however, the
manufacturing cost should be higher due to the harsh lunar
2018-EEF
2018-LEF
2018-LLF
2018-EEC
2018-LEC
2018-LLC
2020-EEF
2020-LEF
2020-LLF
2020-EEC
2020-LEC
2020-LLC
2022-EEF
2022-LEF
2022-LLF
2022-EEC
2022-LEC
2022-LLC
2025-EEF
2025-LEF
2025-LLF
2025-EEC
2025-LEC
2025-LLC
2027-EEF
2027-LEF
2027-LLF
2027-EEC
2027-LEC
2027-LLC
2029-EEF
2029-LEF
2029-LLF
2029-EEC
2029-LEC
2029-LLC
DV (km/sec)
0 5 10 15 20
EEF - Earth- Ear_ - Ful Vehlctl
LEF - Lur',ar- Earth - Full Vehicle
LLF - Lunar - Lunar - Full Vehicle
EEC - Earth- Earlh - Caplule
LEC - Lunar - Earth - C.ap_le
LLC - Lunar - Lunar - Capeule
environment. This paper has assumed that LOX is present in both
LEO and lunar orbit without looking at the manufacturing and
transportation cost. But a relative cost ratio of lunar LOX to LEO
LOX can be assigned, ms shown in Fig. 6. For example, if you .save
400,000 kg of total propellant in LEO by going to the Moon and
picking up 800,000 kg of lunar LOX, then the relative cost ratio
of the lunar LOX is 0.5. In cases where the additional LOX and
H 2 required in LEO to push the MTV to lunar orbit is greater
than the LEO savings incurred by the um of lunar IL_)X, it is not
clear there is a savings and theme missions are assigned a _due
of O in Fig. 6. The results to date are inconclusive as to the viability
of lunar LOX for um in direct manned Mars missions. Additional
study in the areas of propellant production on the Moon and lunar
surface to orbit trans'lx)rtation systems must be conducted before
it can be determined if relative cost ratios such as those shown
in Fig. 6 are technologically feasible. Ks the definition of the lunar
base progresses and Mars mission planning matures, lunar IL)X
propellant production should be analyzed in the framework of
total program life cycle cost.
0.0
Mission
2018-LEF
2018-LLF
2018-LEC
2018-LLC
2020-LEF
2020-LLF
2020-LEC
2020-LLC
2022-LEF
2022-LLF
2022-LEC
2022-LLC
2025-LEF
2025-LLF
2025-LEC
2025-LLC
2027-LEF
2027-LLF
2027-LEC
2027-LLC
2029-LEF
2029-LLF
2029-LEC
2029-LLC
Ratio of
Earth Launch Savings
divided by
Lunar LOX Requirements
0.2 0.4 0.6 ).8 1.0
)
1.2
i
I
EEF - Earth - Earth - Full Vehicle
LEF - Lunar - Earth - Full Vehicle
LLF - Lunar - Lunar - Full Vehicle
EEC - Earth - Earth - Capsule
LEC - Lunar - Earth - Capsule
LLC - Lunar - Lunar - Capsule
Fig. 5. Delta V requirements for "Later" Program options. Fig. 6. Lunar launch cost relative to Earth launch cost.
458 2nd Conference on Lunar Bases and Space Activities
NOTE IN PROOF
This paper presents results only for "direct" manned Mars
missions, i.e., those without gravity assists and without considering
libration point basing options. Several papers have been published
since this paper was presented that give insight into benefits to
be gained from both gravity assist and from libration point basing
options. These publications are now included in the bibliography
of this paper. All these papers are somewhat qualitative in nature,
and it is evident that more refined total life cycle cost analyses
must be accomplished before one basing option can be designated
as optimum for performing manned Mars missions.
REFERENCES
Babb G. and Stump W. (1986) Use of lunar produced propellants for
manned Mars missions. In NASA Manned Mars Study M002, Vol. 1,
pp. 188-200. NASA, Washington, DC.
Corder B. and Wagner S. (1987) Possible Resource Synergisms Between
Mars Programs and Eartl_Moon Activities. General Dynamics Space
Systems Division, San Diego, California. 22 pp.
Davis H. (1986) A manned Mars mission concept with artificial gravity.
In NASA Manned Mars Study M002, VoL 1, pp. 292-302. NASA,
Washington, DC.
Keaton P. W (1985) A Moon Base/Mars Base Transportation DeOot. Los
Alamos Natl. Lab. Rept. LA-lO552-MS. 14 pp.
BIBLIOGRAPHY
Andrews D. G. and Snow W R. (1981) The supply of lunar oxygen to
low Earth orbit. In Papers Presented at the Princeton Space Manufac-
turing Conference II(, pp. 173-179. AIAA, New York
LANL/NASA (1986) Manned Mars Missions (two-volume technical report
of NASA and Los Alamos National Laboratory studies), NASA M002.
Snow W. R., Kubby J. A., and Dunbar R. S. (1981) A small-scale lunar
launcher for early lunar materials utilization. In Papers Presented at the
Princeton Space Manufacturing Conference I_, pp. 157-171. AIAA,
New York
Wc_.xlcock G.R (1985) T_rtation networks for lunar resources
utilization. In Space Manufacturing 5: Engineering with Lunar and
Asteroidal Materials, pp. 89-96. AIAA, New York
Woodcock G. R. (1986a) Logistics support of lunar bases. IAA_6511, 37th
IAE lnnsbruck, Austri_
Woodcock G.R (1986b) Economic potentials for extraterrestrial
resources utilization. 1AA_6-451, 37th IAF, Innsbruck, Austria.
Woodcock G. IL (1988a) Modular evolutionary system architectures for
exploration missions. In Papers Presented to the 1988 Symposium on
Lunar Bases and Space Activities of the 21st Century, p. 258. Lunar
and Planetary Institute, Houston.
Woodcock G. IL (1988b) Basing options for Lunar Oxygen for Manned
Mars MtXs/ons. Boeing Aerospace, Huntsville. 14 pp.


Impact of lunar oxygen production on direct manned Mars missions

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